This invention relates to axial flow turbines, whether driven by gas or steam, of the reaction type.
Reaction type turbines have aerofoils with profiles that cause acceleration of the working fluid along at least a leading region of the suction surface of each aerofoil. The flow over that leading region is laminar but, depending upon the Reynolds number of the aerofoil airflow, the boundary layer further downstream may undergo a transition to turbulent flow and/or there may be transition or separation bubbles formed, which can result in large energy losses.
It is known to give aerofoil surfaces in a turbine, in particular in the low pressure output stages of a gas turbine, a transition-promoting configuration to reduce these losses. In U.S. Pat. No. 4,822,249 (Eckardt et al) a continuous spoiler edge is located closely behind the point of maximum surface velocity on the suction surfaces of the blades of a turbine wheel and extending over substantially the entire radial length of the blades. The function of the spoiler is to promote rapid transition from laminar to turbulent boundary layer flow on the suction surfaces without the formation of laminar separation bubbles. In GB 580806 (Griffith) it is proposed to roughen the entire aerofoil suction surface of reaction type blading for compressors and turbines so as to produce a very thin layer of more or less uniformly disturbed flow over that surface without disturbing the main flow beyond the boundary layer.
The topography of a surface with roughness is complex and there is no single definitive measure of roughness. A widely used basic perimeter is used xe2x80x9caverage roughnessxe2x80x9d (Ra), defined as the arithmetic average of the absolute values of the measured profile height deviations of the surface from the surface profile centreline within a given sampling length. This definition is also valid for previously used alternative terms xe2x80x9carithmetic average roughnessxe2x80x9d (AA) and xe2x80x9ccentreline average roughnessxe2x80x9d (CLA). Typical values of Ra for turbomachinery components are 125 microinches (3.2xc3x9710xe2x88x923 mm) for material as cast and 25 microinches (6.3xc3x9710xe2x88x924 mm) for polished components. Thus, in GB 580806, it has been proposed that the required roughening of the suction surface can be achieved by using sand-cast blades which are not given any polishing or smoothing treatment. In U.S. Pat. No. 5,209,644 (Dorman) which also proposes roughening aerofoils in the output turbine section of a gas turbine that operate in the range of exit Reynolds numbers of 80,000 to 200,000, the surface roughness is in the range 120 to 200 AA microinches (3xc3x9710xe2x88x923 to 5xc3x9710xe2x88x923 mm). Here the roughening is applied to both the suction and pressure surfaces over the entire chord and over most of the aerofoil span and is intended to reduce separation of the boundary layer and formation of recirculation zones or bubbles in the boundary layer.
In the case of modern low pressure turbine blading which operates at low Reynolds numbers (e.g. 70,000-250,000) with highly loaded aerofoil sections, the formation of boundary layer separation bubbles towards the rear of the suction surface cannot be avoided. Steady flow design methodology focusses on ensuring that transition occurs within a bubble, causing it to reattach to the surface as a turbulent boundary layer before the aerofoil trailing edge.
FIG. 1 shows the steady flow isentropic velocity distribution in an annular blading row for a conventional aerofoil (the diamonds plot) and for a high-lift aerofoil (the circles plot) having a lift coefficient approximately 20% greater than the conventional aerofoil. The ordinates are normalised velocities, that is to say, the ordinates are given by the ratio of local flow velocity to exit velocity, and the abscissae are chordal distances normalised as a fraction of the aerofoil chord length measured from the leading edge. The upper pair of plots are for the suction surface and the lower pair are for the pressure surface. A feature of the velocity distribution of high-lift aerofoils, such as that shown in FIG. 1 is the continuous acceleration of flow on the suction surface over the region A from the leading edge to a peak velocity point B typically downstream of the geometric throat in the blade row which will be located at C. The suction side boundary layer is laminar all the way to the peak velocity point B. Deceleration begins after the peak velocity point and in steady viscous flow the boundary layer separates shortly after the start of the deceleration, forming a separation bubble which shows as a plateau up to transition point D. Because transition is reached, the separation bubble reattaches before the trailing edge, resulting in a sharp pressure recovery.
The effectiveness of the high-lift aerofoil design relies on the reattachment of the bubble before the trailing edge because an open separation bubble gives very high losses. Reattachment of the bubble should not occur too early, because that allows unwanted growth of a turbulent boundary layer on the final region of the suction surface which also increases losses.
The preceding discussion, and prior art examples referred to above which seek to avoid the formation of separation bubbles, are all based upon a consideration of aerofoils operating in steady flow. However, in the typical turbine the flow is not steady. There is interaction between succeeding aerofoil rows because the wakes from one row will impinge periodically on the aerofoils of the succeeding row.
A comprehensive review of researches on wake passing effects on separation bubbles is given in xe2x80x9cBlade Row Interactions in Low Pressure Turbinesxe2x80x9d, H P Hodson, von Karman Institute Lecture Series 1998-02, Blade Row Interference Effects in Axial Flow Turbomachinery Stages (1998). The Hodson study considers high-lift, low Reynolds number aerofoils which have been developed for low pressure turbines in order to reduce weight and manufacturing costs. As already mentioned, these aerofoils have regions of significant deceleration on their suction surfaces which can result in the formation of substantial separation bubbles in the absence of wake-passing effects.
The interaction between succeeding aerofoil rows in an axial flow turbine is shown schematically in FIG. 2 (Binder et al) which illustrates how the wakes are transmitted and distorted through downstream rows. The wakes W leaving a first rotor row 2 are chopped by the following stator row 4 and the chopped segments Wxe2x80x2 of the original wakes are further distorted in the flow through the stator row. The dashed lines 6 indicate the stator wakes. Relative to the moving aerofoils of the following rotor row 8, the wake segments are arranged in avenues 10, (indicated by the chains of circles) and if the second rotor row 8 has a different number of blades from the preceding rotor row 2, the respective avenues of wake segments will enter the second rotor row 8 at different phases to the blades of the row.
As is discussed in more detail in the Hodson study, turbulent flow appears in turbomachines typically by bypass transition because of the high levels of turbulence that exist. In this process at points within the boundary layer some distance from the leading edge turbulent spots can form and spread downstream and laterally. Immediately following the rear of a turbulent spot a calmed region is formed having laminar-like characteristics with a very full velocity profile, with a trailing edge travelling at about 30% of the freestream velocity. The unsteady flow of passing wakes can initiate this mechanism to have a beneficial influence on profile losses.
The flow pattern at an optimum wake-passing frequency is illustrated in FIG. 3 which is a space-time diagram of the flow over an aerofoil in which the distance along the aerofoil chord from leading edge to trailing edge is given along the abscissa axis and time values (t/xcfx84) given along the ordinate axis have been normalised by the period of wake passing (T) over the aerofoil.
On the leading part of the aerofoil, to the point of peak velocity, the boundary layer is laminar and is sufficiently stable not to be brought to transition by the wakes of a preceding aerofoil row.
In the steady flow condition described with reference to FIG. 1, beyond the peak velocity point the boundary layer will separate. However, turbulence in a passing wake can create a time-dependent transitional flow regime across the span of the aerofoil. This initiates the development of turbulent spots in the boundary layer transitional flow at each wake passing. Two successive turbulent spots are indicated at F and G. The front of the transitional flow travels at about 90% of the local freestream velocity and when it reaches the trailing edge the boundary layer is turbulent. This is at the point H for the turbulent spot initiated at F.
The rear of the transitional flow travels at only about 50% of the local freestream velocity, so that the chordwise extent and duration of the transitional flow increases as the trailing edge is approached. Furthermore, the transitional flow becomes fully turbulent so that each wake passing is associated with turbulent rather than transitional flow.
Behind each turbulent spot is a calmed region with effectively, laminar flow. The rear of this region travels at about 30% of the freestream velocity and continues to the aerofoil trailing edge. As it passes through the fully turbulent boundary layer, this turbulent region significantly reduces the skin friction locally to below the level of the surrounding, unaffected turbulent boundary layer.
With the optimum wake passing frequency of the regime in FIG. 3, the rear of the becalmed region from one wake reaches the aerofoil trailing edge at the same moment as the front of the turbulent spot initiated by the next wake. Thus, the point I represents both the arrival of the rear of the calmed region at the trailing edge and the arrival there of the front of the transitional flow initiated at the succeeding turbulent spot G. With this situation as shown in FIG. 3, minimum losses are generated by the suction surface boundary layer.
At the nominal separation location, once the calmed region has passed, the deceleration causes the laminar boundary layer to separate and the bubble gradually grows. However, it does not have time to develop fully as it continues only until the next turbulent spot, initiated by the next wake, suppresses it.
More recently, very high-lift aerofoil sections have been developed for low pressure turbine blading which are susceptible to greater profile losses if the boundary layer conditions are not controlled. Such aerofoil sections have a lift coefficient of 1.1 and above as compared with a lift coefficient of 1.0 for high-lift sections. It will be understood that these values relate to the designed normal operating condition within a range of conditions which the turbine might experience. In particular, in the case of aeroengines they usually refer to operation at cruise power The lift coefficient "khgr"2 in cascade flow, assuming compressibility of the flow, is defined as       Ψ    2    =            S              C        AX              ⁢          ρ                        P          o          xe2x80x2                -                  p          e                      ⁢          Va      ⁡              (                              V                          w              ,              e                                -                      V                          w              ,              is                                      )            
where the notation has the following meaning:
Superscriptxe2x80x2 Denotes Isentropic Conditions
FIG. 4 shows, in relation to the radially inner and outer end walls Wi,Wo of a turbine passage in which the aerofoil A operates, the axial chord (CAX) between the leading and trailing edges of the aerofoil section at the mean stream surface.
If incompressible conditions can be assumed, the lift coefficient can be represented more simply as
"psgr"2=S2cos2 xcex22[tanxcex22xe2x88x92tanxcex21]CAX
where:
xcex21 is the inlet flow angle
xcex22 is the exit flow angle
The mechanism described above with reference to FIG. 3, in which an optimum wake passing frequency is chosen to reduce losses, is found to be inapplicable to very high-lift aerofoils. The greater deceleration over the rear part of the suction surface gives a larger separation bubble and transition would occur too late for reattachment. The potential profile losses are consequently greater.
There is therefore a need for an alternative solution in order to take advantage of very high-lift aerofoils the use of which could result in cost and weight benefits.
According to the present invention, an axial flow turbine is provided having downstream of a first row of aerofoils of the turbine, at least one further row of aerofoils in which a region of increased roughness is provided on the suction surface of each aerofoil, said region having an upstream boundary substantially between the location of the geometric throat on that surface and a location 75% of the suction surface perimeter from the leading edge and having an extent of at least 3% of said perimeter.
While this arrangement is able to improve the performance of very high lift aerofoils, the scope of the invention is not limited to such aerofoils. Aerofoils with lower lift characteristics may also experience problems of detachment of flow which, if due to delay of transition, can also be treated in accordance with the present invention.
Preferably, aerofoils according to the invention are designed to operate at a Reynolds number in the range of 10,000 to 250,000. As already mentioned, for aeronautical gas turbines, this typically relates to operation in the cruise condition as the predominant mode of operation.
The region of roughness is preferably not located substantially further forward than the peak velocity point of the aerofoil. For example it can have its leading edge in an area extending from the peak velocity point or from a position not substantially more than 5% of the suction surface perimeter downstream of the peak velocity point.
The aerofoils may, at least at one radial end zone, be operating in an end wall turbulent region, in which case, the region of roughness preferably terminates at or before said zone.